1. Field of the Invention
The present invention relates generally to fluid reaction surfaces and more specifically to air cooled turbine airfoils.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a turbine section with a plurality of stages of stationary vanes or nozzles and rotary blades or buckets receive a hot gas flow from a combustor to produce mechanic power by driving the turbine shaft. The efficiency of the engine can be increased by increasing the hot gas flow through the turbine. However, the turbine materials cannot be used above certain known operating temperatures without being damaged.
In order to allow for turbine parts to operate in conditions above the high temperature material properties will allow, complex air cooling passages have been proposed to cool the vane and blades, especially in the hottest section of the turbine, the first stage immediately downstream from the combustor. Since the cooling air used for cooling the airfoils typically comes from the compressor, the engine efficiency is reduced due to the bleed off air used for cooling. Engine efficiency can also be increased by using maximizing the cooling effect of the air and minimizing the amount of cooling air required.
The prior art airfoil in U.S. Pat. No. 4,529,357 issued to Holland on Jul. 16, 1985 and entitled TURBINE BLADES shows an airfoil cross section having a plurality of radial cooling holes within the wall of the airfoil for cooling the airfoil. This type of cooling passage is one of the least efficient for minimizing the amount of cooling air while maximizing the heat transfer from the airfoil wall to the cooling air. The straight path does not promote much turbulent flow within the cooling air which would increase the convective heat transfer to the cooling air.
An improvement in the Holland, invention above is shown in the U.S. Pat. No. 5,702,232 issued to Moore on Dec. 30, 1997 and entitled COOLED AIRFOILS FOR A GAS TURBINE ENGINE shows an airfoil cross section (FIG. 1) with radial impingement channels 17 spaced along the wall 11 of the airfoil 10, with each channel connected by a metering hole 16 to a cooling supply passage, and each channel having a film cooling hole 18. Cooling air supplied through the supply channel 12 flows through the metering hole 16 for impingement cooling within the radial impingement channels 17, and then out to the airfoil surface through the film cooling holes 18. a showerhead arrangement includes a metering hole 13 connecting the cooling supply channel 12 to a leading edge cavity 14 and film cooling discharge holes 15 spaced along the leading edge. The cooling efficiency of the Moore patent is higher than in the Holland patent. However, in the Moore arrangement the spanwise and chordwise cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In the Moore patent, the spiral flow passages are used in a stationary vane. Applicant's invention is for use in a rotary blade. Rotation of the blade will impart a driving force for the cooling fluid if the spiral passages are in the radial direction as opposed to the chordwise direction in Moore. Also, a single pass radial channel flow is not the best method of utilizing cooling air resulting in low convective cooling effectiveness.
U.S. Pat. No. 4,080,095 issued to Stahl on Mar. 21, 1978 entitled COOLED TURBINE VANE shows a cooled vane with cooling channels having a spiral or twisted arrangement like in a corkscrew-like configuration in which water is passed through for cooling the vane. The spiral shaped cooling passages flow in a chordwise direction of the airfoil and not in a radial direction. Also, the spiral shaped passages do not cross one another such that the two flows will mix.
It is an object of the present invention to provide for a cooling air flow circuit in a turbine airfoil that will provide more cooling using less cooling air than the cited prior art references.